1. Field of the Invention
The present invention relates to gas turbine engines, and more particularly to a system for cooling turbine rotor disk posts of gas turbine engines.
2. Related Art
Conventional high bypass ratio turbofan engines typically include a fan, booster, high pressure compressor, combustor, high pressure turbine and low pressure turbine in axial flow relationship. A portion of the air entering the engine passes through the fan, booster and high pressure compressor, being pressurized in succession by each component. The compressed air exiting the high pressure compressor, commonly referred to as the primary or core gas stream, then enters the combustor where the pressurized air is mixed with fuel and burned to provide a high energy gas stream. However, prior to entering the combustor a portion of the primary or core flow is diverted to provide a source of cooling air for various high temperature components, such as those found in the high pressure turbine. After exiting the combustor, the high energy gas stream then expands through the high pressure turbine where energy is extracted to operate the high pressure compressor which is fixedly connected to the high pressure turbine. The primary gas stream then enters the low pressure turbine where it is further expanded, with energy extracted to operate the fan and booster which are fixedly connected to the low pressure turbine. The remainder of the air flow which enters the engine, other than the primary gas stream continuing through the turbines and the cooling air flow, passes through the fan and exits the engine through a system comprising annular ducts and a discharge nozzle, thereby creating a large portion of the engine thrust.
The highest temperatures in the engine are found in the combustor and turbines. For instance, it is not uncommon for the temperature of the primary gas stream to exceed 2400.degree. F. at the entrance to the first stage blade of the high pressure turbine. The demand for larger and more efficient gas turbine engines creates a requirement for increased turbine operating temperatures, with the metallurgical limitations of critical components such as rotor blades and disks in opposition to this requirement. For example, nickel based alloys are commonly used in the manufacture of turbine rotor disks and with such alloys a typical maximum metal temperature may be approximately 1100.degree. F., which is considerably less than the maximum primary gas path temperature. Consequently, there is a continuing need for novel approaches to provide thermal protection for components such as turbine rotor disks.
A turbine rotor disk is an annular component which rotates about the longitudinal axis of the engine and which supports a plurality of blades that extend radially into the primary gas stream. The disk includes a plurality of circumferentially alternating dovetail slots and posts, with each post formed by adjacent slots, disposed about the periphery of the disk. Each disk dovetail slot is adapted to receive a corresponding dovetail portion, also referred to as a "fir tree" portion, of a blade, with the blades axially loaded into the disk. In addition to the dovetail portion, each blade includes a shank portion which extends radially outward, or away from the engine axis of rotation, from the dovetail portion, and a plate-like platform which radially separates the shank portion from an airfoil portion which extends radially outward from the platform. The outer surface of the blade platforms form a portion of the radially inner boundary of the primary gas stream flowpath, with the platform portions of adjacent stationary nozzle segments forming the remainder of the boundary. For instance, in a multi-stage turbine the stage 1 blades would be positioned between a row of stage 1 nozzles upstream of the blades, which provide the required direction of flow into the blades, and a row of stage 2 nozzles downstream of the blades. Due to the exposure of the blade airfoil to the hot gas stream and the metallurgical limitations of the blade material, it is known in the art to include interior cooling passages in the blade airfoil, with these passages typically supplied through an inner surface of the blade dovetail with the previously discussed compressor discharge air. This cooling air may discharge through a variety of film holes, tip cap holes and trailing edge holes in the blade airfoil, as known in the art.
A structure commonly known as a seal body is typically positioned over the top of each disk post in a cavity bounded by the top of the post, the shank portions of adjacent blades and the underside of the platforms of adjacent blades. The seal body includes a forward plate, which, together with windage baffles typically associated with the axially forward nozzles and the forward portion of the blade platform, or forward angel wing of the blade, forms a cavity on the forward side of the disk. Although this forward cavity is typically purged with the aforementioned compressor discharge cooling air, the cavity air temperature is substantially hotter than the cooling air entering the dovetail portion of the blades due to leakage of hot gases from the primary gas stream into the forward cavity. On an average basis the forward cavity temperature may be 100.degree.-150.degree. F. hotter than the blade cooling air, whereas it may be several hundred degrees higher locally due to rotor/stator non-concentricities or manufacturing tolerances causing increased ingestion of flowpath gases locally. Based on the foregoing, it can be seen that the forward cover plates of the seal bodies discourage the forward cavity purge air, as well as any flowpath gases which are ingested into the forward cavity, from flowing axially between adjacent blade shanks over the tops of the disk posts. The disk post temperature is determined by a heat balance which includes: conduction cooling due to contact with the blade along the dovetail interface; forced convection cooling due to any leakage of blade cooling air which flows through spaces between the disk and blade dovetail serrations, rather than flowing through the blade passages; and convection heating due to hot air surrounding the top of the disk post, with the air being a mixture of the forward cavity air and leakage air from the primary gas stream which is ingested between the platforms of adjacent blades. This air mixture surrounding the top of the disk post is substantially hotter than the blade cooling air. Consequently, thermal isolation of the top of the disk post, from this hot air mixture, is an important part of the overall system for ensuring that the temperature of the disk post does not exceed allowable limits.
Various systems have been employed to provide the necessary isolation of the top of the disk post. One prior disk post isolation system includes shields located at the radially inward side of the blade platforms such that each shield spans the gap between platforms of adjacent blades to discourage ingestion of flowpath gases, and further includes cooling holes through the shank portions of the blades which communicate with the blade interior cooling air passages in order to purge the cavities between the shanks of adjacent blades over each disk post. This system has the disadvantage of placing holes in a highly stressed region of the blade with the stress concentrations associated with the holes creating the potential for cracking and premature failure of the blades. This design has a further disadvantage due to the requirement of purging the relatively large cavities formed between shanks of adjacent blades and bounded at an outer end by the blade platforms and at an inner end by the top of the disk post, which results in the use of a relatively high amount of compressor discharge cooling air and the associated engine performance penalty.
A second prior disk post isolation system, which is illustrated in FIGS. 1 and 2 of this application and subsequently discussed in detail, includes a plurality of seal bodies 31. The seal body 31 illustrated in FIGS. 1 and 2 of this application is substantially the same, with the exception of additional structural features which are present for purposes which are not related to the thermal isolation of disk posts, as the seal body disclosed in U.S. Pat. No. 5,201,849, which is assigned to the assignee of the present invention, and which is herein expressly incorporated by reference. Each seal body 31 has a small diffuser hole 142 extending through a forward cover plate 132 of seal body 31. The entrance of the diffuser hole 142 is in flow communication with the forward cavity 134 and the hole exit is in flow communication with a thermal isolation chamber 144 which is positioned over the top 33 of the disk post 20. Chamber 144 has a relatively small volume as compared to the volume of the cavity bounded by the shanks 26 and platforms 28 of adjacent blades 22 and the radially outer surface 33 of one of the disk posts 20. The diffuser hole 142 causes forward cavity air to slowly drift over the top or radially outer surface 33 of the disk post 20 in order to form an insulative layer of air over the disk post 20. With this system care has to be taken to ensure that the forward cavity air does not pass across the top 33 of the disk post 20 at too high a velocity. Unacceptably high velocities can cause the forced convection from the relatively hot forward cavity air passing across the top 33 of the disk post 20 to dominate the disk post heat balance which can actually result in the disk post temperature rising. Consequently, the system is sensitive to manufacturing tolerances regarding the geometry of diffuser hole 142. One way to obviate the disadvantage associated with the diffuser hole 142 is to lower the temperature of the forward cavity air. However, this would be costly in terms of reduced engine performance due to the relatively high amount of cooling air required to completely purge the forward cavity 134 to effect a reduction in temperature of the forward cavity air. A high amount of cooling air would be required due to the relatively large gaps between the blade angel wing 138 and the associated stator windage baffle structure 128 and stage 1 nozzle inner platform 122 which is required to prevent rubs during engine transient conditions, with the gaps creating a path for ingested gases into the forward cavity 134.
In view of the foregoing, prior to this invention a need existed for a cooling system in a rotor assembly of a gas turbine engine to cool the top of rotor disk posts in a cost effective manner without compromising the structural integrity of the rotor blades and without undue sensitivity regarding the geometry of cooling holes employed.